Method of making integrally bladed rotor

ABSTRACT

Disclosed is a method of making an integrally bladed rotor. According to the method, a rotor disk comprising a radially outer rim surface is provided. A portion of the disk outer rim surface is removed, leaving a protrusion on the rotor disk outer rim surface. The disk with material removed is subjected to thermal processing. A blade comprising an airfoil and a base is positioned such that a base surface is in contact with the protrusion, and heat, pressure, and motion are applied between the blade and the disk to friction weld the base surface to the protrusion.

BACKGROUND

This disclosure is related to the field of bladed rotors generally, andmore specifically to integrally bladed rotors.

Bladed rotors such as impellers, blisks, etc. are employed in gasturbines and other machines. The design, construction and materials ofbladed rotors often dictate operating limits for the turbines in whichthey are employed. Extensive efforts have been made over the years todevelop new alloys, new fabrication techniques, and new componentdesigns which permit operation of these rotors at higher operatingtemperatures and/or lead to lighter weight, longer lived components,with all their attendant advantages.

The fan, turbine, and compressor sections of gas turbine engines includeone or more circumferentially extending rows or stages of airfoils,commonly called rotor blades, which are axially spaced between rows orstages of fixed airfoils (stator vanes). The rotor blades are connectedto and extend radially outwardly from a rotor disk. During operation thecentrifugal loads generated by the rotational action of the rotor bladesmust be carried by the rotor disk within acceptable stress limits.

In one type of a conventional bladed rotor assembly, the rotor disk hasa plurality of slots around its radially outer periphery. The blades maycomprise a root, a platform, and an airfoil. The platform has oppositefacing surfaces. The root attaches to the slot in the disk and theairfoil extends radially out from the disk. The slots and the roots havecomplementary shapes, typically either a dove tail or a fir tree. Theroot mates with the slot and the blade extends radially outwardtherefrom. This type of rotor assembly is relatively heavy because theslots are cut through the rim of the disk creating what is called a“dead rim” where the metal between the slot can pull on the disk withwell over 10,000 g's and fir tree or dovetail mating structures likewisedo not contribute to sustaining the disk's centrifugal load and alsopulls with the same 10,000 g load, thereby necessitating that the rotordisk be sufficiently sturdy, and thus heavy, in order to accommodate thestresses resulting from the heavy blade attachment area.

Alternatively, the blades may be secured by bonding or welding, to therotor disc to thereby form an integrally bladed rotor assembly (IBR). Amajor advantage of an integrally bladed rotor assembly is that there isoften no need for an extended blade root or a blade platform. Theairfoil may be secured directly to the radially outer periphery of therotor disk. The absence of an extended root and a blade platform resultsin a blade that is lighter than a conventional blade. A lighter bladeenables the use of a less rigid and lighter rotor disk, in which casethe integrally bladed rotor assembly is overall much lighter than aconventional bladed rotor assembly.

BRIEF DESCRIPTION

Disclosed is a method of making an integrally bladed rotor. The methodcomprises providing a rotor disk comprising a radially outer rimsurface. A portion of the disk outer rim surface is removed, leaving aprotrusion on the rotor disk outer rim surface. The disk with materialremoved is subjected to thermal processing. A blade comprising anairfoil and a base is positioned such that a base surface is in contactwith the protrusion, and heat, pressure, and motion are applied betweenthe blade and the disk to friction weld the base surface to theprotrusion.

In some embodiments, the method further comprises machining thethermally processed disk outer rim surface to a final shape beforepositioning the blade.

In any one or combination of the foregoing embodiments, the protrusioncan extend in a direction between a leading edge of the disk outer rimsurface to a trailing edge of the disk outer rim surface, and removing aportion of the disk outer rim surface can include removing a portion ofthe disk outer rim surface material along each side of the extendingprotrusion.

In any one or combination of the foregoing embodiments, the protrusioncan be positioned corresponding to a chord of the airfoil where theairfoil meets the disk.

In any one or combination of the foregoing embodiments, the thermalprocessing can comprise heating the disk at or above a metal alloysolution temperature and quenching the disk to a temperature below ametal alloy solution temperature.

In any one or combination of the foregoing embodiments, the methodfurther can comprise holding the disk at or above the metal alloysolution temperature before quenching.

In any one or combination of the foregoing embodiments, the quenchingcan reduce temperature at a rate of at least 100° F. per minute.

In any one or combination of the foregoing embodiments, the quenchingcan reduce temperature at a rate of at least 200° F. per minute.

In any one or combination of the foregoing embodiments, the quenchingcan reduce temperature at a rate of at least 300° F. per minute.

In any one or combination of the foregoing embodiments, the thermallyprocessed disk can comprise a microstructure comprising primary alphagrains of less than 10 μm.

In any one or combination of the foregoing embodiments, the airfoil canbe a solid contiguous structure.

In any one or combination of the foregoing embodiments, each of theblade surface and the recessed area can comprise a titanium alloy.

Also disclosed is an integrally bladed rotor made by the method of anyone or combination of the foregoing embodiments.

Also disclosed is a gas turbine engine comprises an integrally bladedrotor made by the method of any one or combination of the foregoingembodiments.

In some embodiments, a gas turbine engine comprises a fan comprising anintegrally bladed rotor made by the method of any one or combination ofthe foregoing embodiments.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, in which like elements arenumbered alike:

FIG. 1 is a schematic depiction of a gas turbine engine cross-section;

FIG. 2 is a schematic depiction of perspective view of an integrallybladed rotor disk outer rim;

FIG. 3 is a schematic depiction of a perspective view of rotor diskduring fabrication of an integrally bladed rotor disk outer rim prior toremoval of material from the disk outer rim; and

FIG. 4 is a schematic depiction of a perspective view of rotor diskduring fabrication of an integrally bladed rotor disk outer rim afterremoval of material from the disk outer rim before thermal processing.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

As mentioned above, a method is disclosed for friction welding a bladeonto rotor to form an integrally bladed rotor. One application for suchan integrally bladed rotor is on a gas turbine engine, for example as abladed fan. Other applications include compressor rotors or turbinerotors. FIG. 1 schematically illustrates a gas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct, while the compressor section 24 drives air along a coreflow path C for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

With reference now to FIG. 2, a perspective view of a portion of anintegrally bladed rotor 100 is schematically depicted. As shown in FIG.2, an integrally bladed rotor 100 comprises a disk 102 and blades 104.The disk 102 has an outer rim 106 having a radially inner surface 107,and a radially outer surface 108. Each of the blades 104 includes anairfoil 116 and a base 114. The blades attach, at their bases 114, tothe radially outer surface of the rim 106 and extend radially outwardtherefrom. It should be understood that the blades 104 are only two of aplurality of such blades attached to the disk 102. The disk furtherincludes an upstream edge 120 and a downstream edge 122, relative to thegas flow path B (FIG. 1). The blades have upstream and downstream edges124, 126. The integrally bladed rotor 100 typically includes a fillet130, i.e., a curved corner having a radius, adjoining the surfaces ofthe airfoil and the rim. The fillet serves to reduce the concentrationof stress where the airfoils attach to the rim. The fillet itself mayexperience high concentrations of stress at its base 132, where thesurface of the fillet is tangent to the disc rim. A plurality of chords134 each represent a chord of the base of an associated one of theairfoils, i.e. a line between the upstream edge 120 of the base of theairfoil and the downstream edge 122 of the base of the airfoil. Theplurality of chords 134 are typically similarly oriented relative to thelongitudinal axis A (FIG. 1). The chords each lie on an associated oneof a plurality of chord lines 136 that collectively represent thedesired positions of the chords relative to the disk rim.

As mentioned above, the blades 104 are welded to the disk 102 by theapplication of heat (e.g., heat generated by friction), pressure, andmotion, such as by linear friction welding (LFW). As further shown inFIG. 2, a protrusion (also referred to as a standup) 138 extendslinearly with a longitudinal axis along a chord line 136 along thesurface of the rim 106 to receive a base 114 of a blade 104. The weldingcan be initiated by positioning a blade base 114 (not shown in FIG. 3for purposes of illustrating the standup 138) aligned with and incontact with the standup 138. Linear friction welding begins by applyingcompressive and oscillatory forces to the blade base 114. Thecompressive forces can be directed roughly perpendicular to the surfaceof the disk rim 106. The oscillatory forces 136 are directed along anaxis roughly parallel to the longitudinal axis of the standup 138. Theseforces effectively place the interface between the base and standupunder high pressure and cause the base to oscillate back and forthrelative to the rim. At the interface, frictional heat is generated andmaterial from each part changes to a plastic state. Some of thismaterial flows out, in the form of flash, from between the parts,resulting in gradual decrease in the thickness of the parts. Eventually,the interface comprises substantially all points on the opposingsurfaces such that the base surface and the standup surface aresubstantially contiguous. When the process is terminated, the remainingplastic state material of each part cools and changes back to solidstate, forming bonds therein and bonding the two parts together. At theconclusion of the processing, the surface of the base and the recessedsurface are substantially contiguous over the area of the recessedsurface, and the weld between the base and the rim is therefore alsosubstantially continuous.

As mentioned above, a portion or portions of the disk outer rim surfaceare removed, leaving a protrusion or a number of protrusions on the diskouter rim surface. An example embodiment of a rotor disk 102 having adisk outer surface 108 before removal of material is schematically shownin a perspective view in FIG. 3. The rotor disk 102 as shown in FIG. 3can be formed from a metal billet such as a titanium alloy in a forgingpress at a forging temperature. In some embodiments, the metal billetand the rotor disk can be formed according to methods disclosed in U.S.patent application Ser. No. 15/702,486 entitled “Process of MakingIntegrally Bladed Rotor”, filed on even date herewith, the disclosure ofwhich is incorporated herein by reference in its entirety. FIG. 4 showsthe rotor disk 102 after material removal, in which material has beenremoved from the disk outer rim surface in areas 108′, leavingprotrusions or standups 138. Material can be removed using variousprocess such as conventional milling, electro-chemical milling, flankmilling, electrode discharge machining, broaching, grinding. In someembodiments, the material removal performed prior to thermal processingcan leave protrusions that are in net shape or near net shape to serveas standups on which to friction weld the blades 104. In someembodiments, the material removal performed prior to thermal processingcan expose new portions of the disk to the effects of the subsequentlyperformed thermal processing while leaving additional material removalto be performed after thermal processing to achieve net shape or nearnet shape. In either case, a final machining step (e.g., using any ofthe above-mentioned material removal techniques) can be used to provideprecision shaping.

A disk rotor with material removed from the outer rim surface such asshown in FIG. 4 can then be subjected to thermal processing. Thermalprocessing generally involves increasing and reducing the temperature ofthe metal according to a protocol. In some embodiments, for example, thethermal processing can include raising the temperature, holding thetemperature at an elevated temperature, and then cooling. In someembodiments, the hold time at elevated temperature (also called a soakor soak time) can be in a range with a lower end of 10 minutes, or 30minutes, or 1 hour, and an upper end of to 2 hours, or 3 hours, or 4hours. These range endpoints can be independently combined to form anumber of different ranges, and each possible combination of rangeendpoints is hereby expressly disclosed. In some embodiments, theelevated temperature can be at or above a solution temperature of themetal. As used herein, the term “solution temperature” means atemperature at which one or more constituents of the metal alloy formpart of a solid solution. In some embodiments such as for titaniumalloys, the elevated temperature can be in a range having a lower end of75° F. below beta field transition temperature (i.e., beta transustemperature), or 70° F. below beta transus temperature, or 65° F. belowbeta transus temperature, and an upper end of 30° F. below beta fieldtransition temperature, or 25° F. below beta transus temperature, or 20°F. below beta transus temperature. These range endpoints can beindependently combined to form a number of different ranges, and eachpossible combination of range endpoints is hereby expressly disclosed.As used herein, the beta field transition or beta transus temperature isthe lowest temperature at which the alloy can exist in a 100% betaphase.

In some embodiments, the cooling can be performed rapidly (also known asquenching). Although this disclosure is not bound by any theory ofoperation, it is believed that a rapid reduction in temperature can leadto the formation of beneficial grain microstructures in the metal. Insome embodiments, temperature reduction occurs at a rate of at least100° F./minute, or at least 200° F./minute, or at least 300° F./. Insome embodiments, the temperature reduction rate can be in a range witha lower end of 300° F./minute, or 200° F./minute, or 300° F., and anupper end of to 700° F./minute, or 850° F./minute, or 1000° F. Theserange endpoints can be independently combined to form a number ofdifferent ranges, and each possible combination of range endpoints ishereby expressly disclosed. In some embodiments, the above cooling ratecan be maintained during a transition from the soak temperature down toa temperature of 1100° F. In some embodiments, a resultant metal grainstructure can include microstructure with fine primary grain size lessthan 10 um with a volume fraction of 15-50%, and a widmanstattensecondary alpha (<1.0 um). In some embodiments, the metal grainstructure can include microstructure with a low volume fraction ofprimary alpha (<30%) grains.

Various materials can be used for the blade base and disk rim. In someembodiments, both the base and the rim comprise nickel alloys, which canbe the same alloy for both parts or different alloys. For example, insome embodiments, the blade can be formed from an alloy such as PWA 1228or equivalent and the disk can be formed from an alloy such as PWA 1215,PWA 1214, or equivalent. In some embodiments, the blade and the base canbe formed from the same material, such as PWA 1215 or equivalent.

In some embodiments, the embodiments described herein can provide atechnical effect promoting metal grain microstructures for a robustintegrally bladed rotor structure. The formation of protrusions orstandups on a disk rim for friction welding attachment of blades canprovide benefits such as ease of fabrication and promoting ejection orremoval of flash from the weld zone. However, it has now been discoveredthat the machining away of metal from the disk rim to form the standupscan remove superior metal with fine grain structure properties whichtends to exist in disk outer diameter areas, exposing coarser grainmaterial that is less resistant to stress from vibration. This stresscan be particularly severe with some designs such as solid (as opposedto hollow) fan blades. Thermal processing of the rotor disk after thematerial removal can provide a technical effect of providing targetsurface(s) on the disk outer rim for blade attachment through frictionwelding while also promoting formation of a region of fine grainstructure metal in the integrally bladed rotor structure.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application. For example, “about”can include a range of ±8% or 5%, or 2% of a given value.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. A method of making an integrally bladed rotor,comprising providing a rotor disk comprising a radially outer rimsurface free from blades; removing a portion of the disk outer rimsurface, leaving a protrusion on the rotor disk outer rim surface;subjecting the disk including said protrusion to thermal processing;positioning a blade comprising an airfoil and a base such that a basesurface is in contact with the thermally processed protrusion; andapplying heat, pressure, and motion between the blade and the disk tofriction weld the base surface to the protrusion.
 2. The method of claim1, further comprising machining the thermally processed disk outer rimsurface to a final shape before positioning the blade.
 3. The method ofclaim 1, wherein the protrusion extends in a direction between a leadingedge of the disk outer rim surface to a trailing edge of the disk outerrim surface, and removing a portion of the disk outer rim surfaceincludes removing a portion of the disk outer rim surface material alongeach side of the extending protrusion.
 4. The method of claim 3, whereinthe protrusion is positioned corresponding to a chord of the airfoilwhere the airfoil meets the disk.
 5. The method of claim 1, wherein thedisk comprises a metal alloy, and the thermal processing comprisesheating the disk at or above a solution temperature of the metal alloyand quenching the disk to a temperature below the solution temperatureof the metal alloy.
 6. The method of claim 5, further comprising holdingthe disk at or above the solution temperature of the metal alloy beforequenching.
 7. The method of claim 5, wherein the quenching reducestemperature at a rate of at least 100° F. per minute.
 8. The method ofclaim 5, wherein the quenching reduces temperature at a rate of at least200° F. per minute.
 9. The method of claim 5, wherein the quenchingreduces temperature at a rate of at least 300° F. per minute.
 10. Themethod of claim 1, wherein the thermally processed disk comprises amicrostructure comprising primary alpha grains of less than 10 μm. 11.The method of claim 1, wherein the airfoil is a solid contiguousstructure.
 12. The method of claim 1, wherein each of the blade and thedisk comprises a titanium alloy.